Al-si-mg aluminum alloy aircraft structural component production method

ABSTRACT

An aircraft structural component, particularly a fuselage component, production method using rolled, extruded or forged products made of aluminum alloy with the following composition (% by weight):  
     Si: 0.7-1.3; Mg: 0.6-1.1; Cu: 0.5-1.1; Mn: 0.3-0.8; Zn&lt;1; Fe&lt;0.30; Zr&lt;0.20; Cr&lt;0.25; other elements &lt;0.05 each and &lt;0.15 in total; remainder aluminum. The product is treated by steps of:  
     solution heat treatment of the product between 540 and 570° C.;  
     quenching the solution heat treated product;  
     producing the structural component by forming of the product, and welding if required;  
     ageing of the structural component, in one or more stages, for which the total equivalent time at 175° C. expressed in hours is between (−160+57γ) and (−184+69γ), γ being the sum of Si+2Mg+2Cu contents in % by weight. The invention results in an improvement in tolerance to damage with no loss of other operating parameters.

FIELD OF THE INVENTION

[0001] The invention relates to the field of rolled, extruded or forgedproducts made of series 6000 Al—Si—Mg aluminum alloy according toAluminum Association references, intended to produce aircraft structuralcomponents, particularly fuselage components.

[0002] 2. Description of Related Art

[0003] Commercial aircraft fuselages are mostly produced using 2024 T3or T351 alloy sheets, clad on both sides with a low content aluminumalloy, for example a 1050 or 1070 alloy, in order to improve corrosionresistance. The thickness of the cladding may represent, depending onthe thickness of the core sheet, between 2 and 12% of the totalthickness.

[0004] Several years ago, it was proposed to use series 6000 Al—Si—Mgalloys for fuselage panels instead of 2024 alloys or similar alloys.These alloys, also heat-treated, offer good mechanical characteristicswhen treated, a high modulus of elasticity and a lower density than2024. The alloys are easier to weld, enabling a reduction in the numberof riveted assemblies, which are a source of additional cost, and alsosites in which stress is concentrated and corrosion initiated.

[0005] The patent U.S. Pat. No. 4,589,932 (Alcoa) discloses the use, foraircraft structural components, of an alloy subsequently registeredunder the reference 6013, with the following composition (% by weight):

[0006] Si: 0.4-1.2 Mg: 0.5-1.3 Cu: 0.6-1.1 Mn: 0.1-1 Fe<0.6

[0007] The patent EP 0 173 632, filed by the applicant, discloses analloy, subsequently registered under the reference 6056, with thefollowing composition:

[0008] Si: 0.9-1.2 Mg: 0.7-1.1 Cu: 0.3-1.1 Mn: 0.25-0.75 Zn: 0.1-0.7 Zr:0.07-0.2 Fe<0.3

[0009] The patent EP 0 787 217, also filed by the applicant, relates toa specific ageing treatment, resulting in a T78 temper, for a 6056 typealloy, so as to desensitize it to intercrystalline corrosion, and thusenable its use without cladding for aircraft fuselages. This ageing isdefined by a total duration, measured in equivalent time at 175° C.,between 30 and 300 hours, and preferentially between 70 and 120 hours.This development was the subject of a presentation by R. Dif, D. Béchet,T. Warner and H. Ribes: “6056 T78: A corrosion resistant copper-rich6xxx alloy for aerospace applications” at the ICAA-6 congress (July1998) in Toyohashi (Japan) and published in the Congress Proceedings,pages 1991-1996.

[0010] The parts are preferentially shaped in the T4 temper, wherein the6056 alloy shows excellent forming properties. The ageing is performedon shaped and possibly welded parts. The use of 6056-T78 results incomplete desensitisation to intercrystalline corrosion of the weldedjoin or the base product and in static mechanical characteristicsequivalent to those of clad T3 or T351 2024. However, it seemeddesirable to improve the results obtained in terms of tolerance todamage, while retaining the static mechanical properties anddesensitization to intercrystalline corrosion.

SUMMARY OF THE INVENTION

[0011] The invention relates to an aircraft structural componentproduction method using rolled, extruded or forged products made ofaluminum alloy comprising:

[0012] the casting of a blank with the following composition (% byweight):

[0013] Si: 0.7-1.3 Mg: 0.6-1.1 Cu: 0.5-1.1 Mn: 0.3-0.8 Zn<1 Fe<0.30Zr<0.20 Cr<0.25 other elements <0.05 each and <0.15 in total, theremainder being aluminum,

[0014] hot, and possibly cold, transformation of said blank to produce aproduct,

[0015] solution heat treatment of the product between 540 and 570° C.,

[0016] quenching of said product,

[0017] production of the structural component by forming of the product,and welding if required,

[0018] ageing of the structural component, in one or more stages, forwhich the total equivalent time at 175° C. expressed in hours is between(−160+57γ) and (−184+69γ), γ being the sum of the Si+2Mg+2Cu contents in% by weight.

[0019] The invention also relates to an aircraft structural componentproduction method, wherein the composition of the products belongs to apreferential composition range (% by weight):

[0020] Si: 0.7-1.1 Mg: 0.6-0.9 Cu: 0.5-0.7 Mn: 0.3-0.8 Zr<0.2 Fe<0.2Zn<0.5 Cr<0.25 Mg/Si<1, Si+2 Mg: 2-2.6 other elements <0.05 each and<0.15 in total, the remainder being aluminum,

[0021] and the ageing time is between 40 and 65 hours of totalequivalent time at 175° C.

[0022] It also relates to an aircraft fuselage component produced usingproducts with the preferential composition given above.

BRIEF DESCRIPTION OF THE DRAWINGS

[0023]FIG. 1 represents, in the form of Wöhler curves, the fatigueservice life of T6 and T78 temper samples according to Example 1, beforeand after prolonged exposure in a marine environment.

[0024]FIG. 2 represents the results of intercrystalline corrosion testsas a function of the yield strength in the TL direction in the T4 temperfor the samples in Examples 6 and 7.

DETAILED DESCRIPTION OF THE INVENTION

[0025] The invention is based on the observation that within thecomposition and ageing range disclosed in the patent EP 0 787 217, thereis a restricted range linking the major elements of the composition (Si,Mg and Cu) and the total equivalent ageing time at 175° C., as thisparameter is defined in EP 0 787 217; with this range, in relation tothe results given in the examples of this European patent, animprovement in the static mechanical characteristics and tolerance todamage is obtained, with no adverse effect on sensitivity tointercrystalline corrosion. It is thus possible to associate with eachalloy composition a factor γ equal to the sum of the Si+2Mg+2Cu contents(in % by weight) and with said factor γ a period of equivalent ageingtime at 175° C. between (in hours) (−160+57γ) and (−184+69γ) andpreferentially between (−150+57γ) and (−184+69γ).

[0026] More specifically, the inventors revealed that by unloading thealloy in relation to compositions of the examples in the Europeanpatent, i.e. by positioning at the lower end of the content ranges forthese 3 elements, while ensuring that these elements are put in solutionas completely as possible, the alloy became less sensitive tointercrystalline corrosion at given over-ageing and that, as a result,it was possible to desensitize it with a lower level over-ageing.

[0027] In this way, in the preferential composition range mentionedabove, with particularly Cu<0.7% and Si+2Mg<2.6%, the equivalent ageingtime at 175° C. to reach the T78 temper with total desensitization isbetween 40 and 65 hours, i.e. below the preferential range (70 to 120hours) indicated in the patent EP 0 787 217. However, to obtain asufficient mechanical resistance, it is necessary to maintain Cu>0.5%and Si+2Mg>2.0 and preferentially >2.3%.

[0028] In this preferential composition range, associated with T78ageing for an equivalent time at 175° C. between 40 and 65 hours, it ispossible to obtain, in addition to complete desensitization tointercrystalline corrosion, the following level of properties in termsof static mechanical characteristics, toughness and crack growth:

[0029] a yield strength R_(0.2) (TL direction)>330 MPa, an ultimatetensile strength R_(m) (TL direction)>360 MPa and an elongation A (TLdirection)>8%.

[0030] a plane strain fracture toughness, measured in the T-L direction,according to the ASTM E561 standard, such that at least one of thefollowing properties is verified:

[0031] K_(R) (Δa=20 mm)>90 MPa{square root}m

[0032] K_(R) (Δa=40 mm)>115 MPa{square root}m

[0033] K_(c0)>80 MPa{square root}m

[0034] K_(c)>110 MPa{square root}m

[0035] The measurements are made on a CCT test specimen of width W=760mm and initial cracking length 2a₀=253 mm. The test makes it possible todefine the R curve of the material, giving the tear strength K_(R) as afunction of the crack extension Δa. Using this curve, it is thenpossible to calculate, according to the procedures indicated by L.Schwarmann in Aluminium, 1991, vol. 67, No. 5, p. 479, the apparentfracture toughness K_(c0) and effective fracture toughness K_(c) whichcorrespond to the break of a virtual CCT type test specimen of widthW=400 mm and initial crack length 2a₀=133 mm.

[0036] a fracture toughness in the L-T direction, measured under thesame conditions as that in the T-L direction, such that at least one ofthe following properties is verified:

[0037] K_(c0)>90 MPa{square root}m

[0038] K_(c)>130 MPa{square root}m

[0039] a crack growth rate da/dn, measured in the T-L directionaccording the ASTM E647 standard for R=0.1 on a CCT type test specimenof width W=160 mm, less than:

[0040] 2 10⁻³ mm/cycle for ΔK=20 MPa{square root}m

[0041] 4 10⁻³ mm/cycle for ΔK=25 MPa{square root}m

[0042] 8 10⁻³ mm/cycle for ΔK=30 MPa{square root}m

[0043] Finally, in this particular T78 temper, a lower drop in thefatigue resistance after prolonged exposure in a corrosive environmentis observed in relation to the T6 temper.

[0044] This set of properties, associated with the fact that the alloycan be welded, makes it particularly suitable for the production ofaircraft structural components, particularly fuselage components.

[0045] It is also possible to use the alloy, in the preferentialcomposition according to the invention, in the T6 temper. The level ofproperties obtained in said T6 temper with the preferential compositionaccording to the invention, in terms of static mechanicalcharacteristics, fracture toughness and crack growth rate is as follows:

[0046] a yield strength R_(0.2) (TL direction)>350 MPa, an ultimatetensile strength R_(m) (TL direction)>380 MPa and an elongation A (TLdirection)>6%.

[0047] a fracture toughness, measured under the same conditions as forthe T78 temper mentioned above, such that at least one of the followingproperties is verified:

[0048] K_(R) (Δa=20 mm)>95 MPa{square root}m

[0049] K_(R) (Δa=40 mm)>120 MPa{square root}m

[0050] K_(c0)>85 MPa{square root}m

[0051] K_(c)>115 MPa{square root}m

[0052] a fracture toughness measured in the L-T direction under the sameconditions, such that at least one of the following properties isverified:

[0053] K_(c0)>100 MPa{square root}m

[0054] K_(c)>150 MPa{square root}m

[0055] a crack growth rate da/dn, measured under the same conditions asfor the T78 temper, less than:

[0056] 2 10⁻³ mm/cycle for ΔK=20 MPa{square root}m

[0057] 4 10⁻³ mm/cycle for ΔK=25 MPa{square root}m

[0058] 8 10⁻³ mm/cycle for ΔK=30 MPa{square root}m

[0059] This set of properties, associated with the fact that the alloycan be welded, makes the product particularly suitable for theproduction of aircraft fuselage components.

[0060] The production method according to the invention comprises thecasting of a blank made of the composition mentioned, said blank being aplate for rolled products, a slug for extruded products or a forgingblock for forged products. The blank is scalped and then heated beforehot transformation by rolling, extrusion or forging, and possiblyundergoes cold transformation. After cutting in the required format, thesemi-finished product obtained undergoes a heat treatment at atemperature between 540 and 570° C., quenched, generally in cold water,and finished, the purpose of said final step essentially being to absorbthe deformations of the semi-finished product after quenching. Theproduct is most frequently supplied in this T4 temper to shape thestructural component and for welding if required. The formed, and ifapplicable welded, component then undergoes the ageing treatmentaccording to the invention.

[0061] The applicant observed that it is advantageous to add, beforescalping, a homogenization step at a temperature between 540 and 570°.Said homogenization may comprise a single stage, or two stages, thesecond stage being at a higher temperature than the first. It helpsimprove the forming properties of the product in the T4 temper andreduce the grain size, leading to a decrease in the roughness of themetal when it undergoes chemical machining. Excessive roughness mayinduce initial micro-cracking due to fatigue.

[0062] In addition, the tests demonstrated that the desensitization tointercrystalline corrosion increases in effectiveness with the level ofcold-working in the T4 temper. This cold-working may be a result offinishing operations such as straightening or planing with rollers ortraction for sheets and traction or drawing for profiles. It may also bethe result of part forming operations by rolling, drawing-forming,embossing, flow turning or folding. Said cold-working, of at least 1%,and preferably of at least 2% permanent elongation, may be relativelysignificant, for example up to 10%, or even up to 15% permanentelongation; indeed, it is observed, surprisingly, that significantcold-working, although it accelerates the ageing kinetics, does notreduce the yield strength in the T78 temper with reference to the samenon-cold-worked product.

[0063] This possibility to use significant cold-working improving theresistance to intercrystalline corrosion is particularly useful if, asis frequently the case for aircraft fuselage production, thin sheets andprofiles must be assembled. Indeed, the applicant observed thatdesensitization to intercrystalline corrosion is more difficult to carryout on profiles than on sheets, probably due to the difference in theirmetallurgic structure. If the sheets and profiles are formed separately,and then welded before ageing, this is liable to induce a difference incorrosion resistance between the sections produced from profiles andthose produced from sheets. To remedy this disadvantage, rather thanchoose a very high level of ageing to desensitize the profiles, whichwould induce a significant loss of mechanical resistance, it ispreferable to retain the T78 ageing adapted to the desensitization ofthe sheets and to subject the profiles to an additional cold-workingstep to bring their resistance to intercrystalline corrosion to the samelevel as those of thin sheets.

EXAMPLES Example 1

[0064] A plate of the composition (% by weight) corresponding to Example3 of the patent EP 0 787 217 was cast, i.e.: Si: 0.92 Mg: 0.86 Cu: 0.87Mn: 0.55 Fe: 0.19 Zn: 0.15 Zr: 0.10 i.e. Mg/Si=0.93 and Si+2Mg=2.64.

[0065] The plate was heated at 530° C., scalped, hot and then coldrolled to a thickness of 3.2 mm. Samples of the sheet obtained weresubjected to a solution heat treatment at 550° C., quenched in water,finished and subjected to ageing. In some cases, the ageing lasted 8hours at 175° C. to obtain the T6 temper, i.e. the temper correspondingto maximum mechanical resistance; in other cases, it lasted 6 hours at175° C. and then 2 hours at 220° C., or an equivalent time at 175° C. of95 hours, to obtain the T78 temper, as described in Example 3 of thepatent EP 0 787 217.

[0066] The mechanical characteristics were measured in the TL direction,i.e. the tensile strength Rm (in MPa), the conventional yield strengthat 0.2% elongation R_(0.2) (in MPa) and the fracture elongation A (in%), along with the sensitivity to intercrystalline corrosion ICaccording to the US Army standard MIL-H-6088. Complete desensitizationis defined as the absence of corrosion ramifications over 5 μm long. Theresults are given in Table 1. TABLE 1 IC Temper R_(0.2) (TL) R_(m) (TL)A (TL) sensitivity T6 364 408 7 Yes T78 304 343 8 No

[0067] For the T78 temper, the fracture toughness was also measuredusing the R curve method, according to the ASTM E 561 standard. Thetest, performed on a CCT type test specimen of width W=760 mm andcentral cracking length 2a₀=253 mm, is used to deduce the curve linkingthe tear strength K_(R) to the increase in cracking Δa. For the T-Ldirection, the value of K_(R) for increases in cracking Δa=20 mm andΔa=40 mm is given in Table 2.

[0068] The R curve is also used, for example using L. Schwarmann'smethod mentioned above, to determine by calculation the plane straintoughnesses K_(c0) (apparent toughness) and K_(c) (effective toughness),in MPa{square root}m, which correspond to the critical stress intensityfactors for a CCT test specimen, with a width W=400 mm and initialcracking length 2a₀=133 mm. The results in the T-L and L-T directionsare also given in Table 2: TABLE 2 K_(R) (T-L) K_(R) (T-L) K_(c0) K_(c)K_(c0) K_(c) Temper Δa = 20 mm Δa = 40 mm (T-L) (T-L) (L-T) (L-T) T7889.5 107.5 75.2 105.9 88.8 137.8

[0069] The fatigue crack growth was also measured in the T78 temper inthe T-L direction (in mm/cycle) for R=0.1 (ratio between minimal andmaximal stress) and for different values of ΔK (in MPa{square root}m)according to the ASTM E 647 standard. The results, obtained on CCT typetest specimens of width W=160 mm, are given in Table 3: TABLE 3 TemperΔK = 20 MPa{square root}m ΔK = 25 MPa{square root}m ΔK = 30 MPa{squareroot}m T78 10⁻³ 3 10⁻³ 6.3 10⁻³

Example 2

[0070] A plate of a composition included in the preferential compositionof the present invention was cast: Si=0.93 Mg=0.75 Cu=0.60 Mn=0.63Fe=0.10 Zn=0.16 which corresponds to Mg/Si=0.81 and Si+2Mg=2.43.

[0071] The plate was transformed under the same conditions as in Example1, except in terms of the ageing in the T78 temper. Part of the samplesunderwent ageing for 6 hours at 175° C. followed by 5 hours at 210° C.,or a total equivalent time at 175° C. of 105 hours, according to thepreferential disclosure in the patent EP 0787217. Another part underwentageing for 6 hours at 175° C. followed by 13 hours at 190° C., or atotal equivalent time at 175° C. of 55 hours, according to the presentinvention. The same measurements as in Example 1 were made for the T6and T78 tempers at 105 hours and 55 hours. The results are given inTables 4, 5 and 6. TABLE 4 IC Temper R_(0.2) (TL) R_(m) (TL) A (TL)sensitivity T6 360 397 7.5 Yes T78 305 337 10.5  No (105 hrs) T78 339367 9.2 No (55 hrs)

[0072] It is observed that ageing for 55 hours equivalent time improvesmechanical resistance significantly in relation to that for 105 hoursequivalent time, while showing the same desensitization tointercrystalline corrosion. TABLE 5 K_(R) (T-L) K_(R) (T-L) K_(c0) K_(c)K_(c0) K_(c) Temper Δa = 20 mm Δa = 40 mm (T-L) (T-L) (L-T) (L-T) T6101.1  126.2 87.9 121.7 104.4 155.1 T78 94.4 119.6 83.1 117.5  91.6137.9 105 hrs T78 96.5 125   86.9 125.7 55 hrs

[0073] It is observed firstly that, with the same ageing, the variationin composition between Example 1 and Example 2 results in an improvementin fracture toughness, irrespective of the measurement parameter usedand secondly that, with the same composition, the ageing for 55 hoursequivalent time also improves the toughness. TABLE 6 Temper ΔK = 20MPa{square root}m ΔK = 25 MPa{square root}m ΔK = 30 MPa{square root}m T61.2 10⁻³ 3 10⁻³ 5 10⁻³ T78 (105 hrs)    10⁻³ 2 10⁻³ 4 10⁻³ T78 (55 hrs)1.2 10⁻³ 3 10⁻³ 5 10⁻³

[0074] It is observed that with the ageing and preferential compositionaccording to the invention, there is no degradation of da/dn between T6and T78.

[0075] On the same sheets in the T6 and T78 temper, fatigue specimenblanks were removed and exposed for one year to a marine environment onthe Mediterranean coast. After machining, the test specimens, showing astrain concentration factor of almost 1, underwent fatigue-endurancetests, to determine the number of fracture cycles, at different levelsof strain and a frequency of 30 Hz, for a load ratio R=0.1. The resultsare represented in FIG. 1 in the form of Wöhler curves, both on thenon-corroded material (solid lines) and on the corroded test specimens(individual dots).

[0076] These results demonstrate the advantage of the T78 treatment withreference to the T6 treatment in terms of the drop in fatigue resistanceafter exposure to corrosion.

Example 3

[0077] Three plates made of alloys A, B and C were cast, for which thecompositions (by weight %) within the preferential composition rangeaccording to the invention and the final rolling thicknesses e, aregiven in Table 7: TABLE 7 Alloy E (mm) Si Mg Cu Mn Fe Zn Si + 2 Mg A1.4-3.2 0.93 0.75 0.60 0.63 0.10 0.16 2.43 B   4-8 0.91 0.76 0.64 0.590.13 0.17 2.43 C 4.5-6   0.94 0.80 0.64 0.56 0.10 0.13 2.54

[0078] The plates were transformed in the same way as those in the aboveexamples up to the ageing step, apart from the fact that, forthicknesses greater than or equal to 4.5 mm, given in Table 7, no coldrolling was carried out. The same ageing for 6 hours at 175° C.+13 hoursat 190° C., or a total equivalent time at 175° C. of 55 hours, wasperformed for all the samples. The same measurements as in the aboveexamples were made: static mechanical characteristics (TL direction)R_(0.2) (in MPa), Rm (in MPa) and A (in %), sensitivity tointercrystalline corrosion, fracture toughness (T-L direction) and crackgrowth rate (T-L direction). The results are given in Tables 8, 9 and10. TABLE 8 IC Alloy - th R_(0.2) (TL) R_(m) (TL) A (TL) sensitivity A1.4 mm 337 363 8.3 No A 3.2 mm 339 367 9.2 No B 4 mm 340 369 9.1 No B 8mm 345 371 8.9 No C 4.5 mm 337 367 9.4 No C 6 mm 351 379 9.4 No

[0079] TABLE 9 K_(R) (T-L) K_(R) (T-L) Alloy - th Δa = 20 mm Δa = 40 mmK_(c0) (T-L) K_(c) (T-L) A 1.4 mm 90   122.5 85.5 129.7 A 3.2 mm 95.5125   86.9 125.7 B 8 mm 110   134   93.8 126.1 C 4.5 mm 98.5 121.5 84.9114.7

[0080] TABLE 10 Alloy - th ΔK = 20 MPa{square root}m ΔK = 25 MPa{squareroot}m ΔK = 30 MPa{square root}m A 1.4 mm 1.3 10⁻³ 2.5 10⁻³ 5.2 10⁻³ A3.2 mm 1.1 10⁻³   3 10⁻³ 4.8 10⁻³ B 8 mm   8 10⁻⁴ 2.3 10⁻³ 4.1 10⁻³ C4.5 mm 1.1 10⁻³ 2.8 10⁻³ 4.3 10⁻³

[0081] It is observed that, for all the thicknesses, irrespective ofwhether cold rolling was performed or not, the values measured for thestatic mechanical characteristics and toughnesses are greater than theminimum values given above for the T78 temper, and the crack growthsda/dn are less than the maximum values given above for the same temper.

Example 4

[0082] An alloy with the following composition (% by weight) was cast:Si=1.01 Mg=0.71 Cu=0.67 Mn=0.55 Fe=0.14 Zn=0.15 remainder aluminum.

[0083] A first plate of this alloy was subjected to the productionprocedure A comprising the following steps: homogenization for 4 hoursat 540° C.+24 hours at 565° C., scalping, heating at 530° C., hotrolling of a strip up to 4.5 mm, conversion of strip into sheets,solution heat treatment in a furnace for 40 min at 550° C. in air, waterquenching, finishing, T6 ageing for 8 hours at 175° C.

[0084] A second plate underwent production procedure B comprising thesame steps except for the preliminary homogenization. The grain size(thickness e and length l) on the surface and mid-thickness of the sheetwas measured in the T4 temper (before ageing) by optical microscopy on aground section, along with the distribution of the Al—Mn—Si dispersoidsin electron microscopy in transmission. This distribution is evaluatedwith the parameter ECD (Equivalent Circle Diameter)={square root}4A/πwherein A in the area of the phases observed on the microscopic section.To characterize the formability, the parameter LDH (Limit Dome Height)is used. This parameter is defined in the patent application EP 1045043filed by the applicant. The results are given in Table 11: TABLE 11 egrain 1 grain e grain 1 grain surface surface mid-th. mid-th. ECD LDHProcedure (μm) (μm) (μm) (μm) (nm) (mm) A 27 143 23 140 271 92 B 40 31630 320 108 73

[0085] It is observed that, in the T4 temper, i.e. in the temper inwhich sheets are most frequently delivered to the aeronauticmanufacturer for forming, ageing, procedure A with homogenizationresults in a smaller grain size and therefore lower roughness afterchemical machining and improved formability.

[0086] The static mechanical characteristics R_(0.2) (in MPa), Rm (inMPa) and A (in %) in the L and TL directions in the T6 temper were alsocompared for both procedures. The results are given in Table 12: TABLE12 R_(0.2) Procedure (TL) R_(m) (TL) A (TL) R_(0.2) (L) R_(m) (L) A (L)A 361 390 11.3 374 386 12.0 B 359 389 10.5 367 386 12.7

[0087] It is possible to conclude that homogenization does not have asignificant effect on the mechanical characteristics in the T6 temper.

Example 5

[0088] An alloy with the following composition (% by weight) was cast:Si=0.82 Mg=0.68 Cu=0.55 Mn=0.57 Fe=0.13 Zn=0.14 remainder aluminum.

[0089] This plate alloy was subjected to the following productionprocedure: homogenization for 4 hours at 540° C.+24 hours at 565° C.,scalping, heating at 530° C., hot rolling of a strip up to 5 mm,conversion of strip into sheets, solution heat treatment in a furnacefor 40 min at 550° C. in air, water quenching, finishing, T78 ageing for6 hours at 175° C.+13 hours at 190° C. (or 55 hours of equivalent timeat 175° C.).

[0090] The static mechanical characteristics R_(0.2), R_(m) (in MPa) andA (in %) in the TL direction were measured in said T78 temper, alongwith the fracture toughness in the T-L direction (in MPa{square root}m),the crack growth rate in the T-L direction and the sensitivity tointercrystalline corrosion in the same way as for Examples 1, 2 and 3.The results are given in Tables 13, 14 and 15: TABLE 13 IC ProcedureR_(0.2) (TL) R_(m) (TL) A (TL) sensitivity Homog. + 337 359 11 No T78

[0091] TABLE 14 K_(R) (T-L) K_(R) (T-L) Temper Δa = 20 mm Δa = 40 mmK_(c0) (T-L) K_(c) (T-L) Homog. + 115 142 98.8 136 T78

[0092] TABLE 15 Temper ΔK = 20 MPa{square root}m ΔK = 25 MPa{squareroot}m ΔK = 30 MPa{square root}m Homog. + T78 1.1 10⁻³ 2.1 10⁻³ 4.0 10⁻³

[0093] If these results are compared to those in Table 4 of Example 2,it is noted that, also in the T78 temper, homogenization does not have asignificant effect on mechanical characteristics, the crack growth rateor the sensitivity to intercrystalline corrosion, but seems to increasethe fracture toughness measured by the R curve.

Example 6

[0094] Samples were taken from the sheets of Example 3 and Example 5 atdifferent thicknesses and with different types of finishing, comprisingat least one of the straightening D, roller planning P or tractionplaning T operations. In each case, the yield strength R_(0.2) in the TLdirection (in MPa) in the T4 and T78 temper was measured, along with thesensitivity to intercrystalline corrosion in the T78 temper. Thiscorrosion was qualified as “slight” when it induces pitting with shortintergranular ramifications. The results are given in Table 16: TABLE 16Alloy R_(0.2) (TL) ex. e (mm) Finish T4 R_(0.2) T78 IC sens. 3 A 1.4 D +P + T 218 337 No 3 A 3.2 D + P + T 215 339 No 3 B 4 D + P + T 218 340 No3 B 8 T 181 345 Slight 3 C 4.5 D + P + T 203 337 No 3 C 6 T 198 351Slight 5 2.2 D + P + T 179 340 Yes 5 2.2 D + P + T 211 336 Slight 5 2.5D + P + T 224 332 No 5 2.5 D + P + T 200 330 Slight 5 3.2 D + P + T 245326 No 5 5 P + T 218 337 No

[0095] The results given in FIG. 2 show, for a given composition, aclear correlation between the resistance to intercrystalline corrosionin the T78 temper and the yield strength in the T4 temper.

Example 7

[0096] Using the sheet corresponding to the seventh row of Table 14(composition according to Example 5, thickness 2.2 mm), differentfinishing operations were carried out in the laboratory in the T4 temperconsisting of controlled traction to 3.2% permanent elongation and coldrolling to different levels of permanent elongation between 2.6 and8.7%. The samples obtained in this way were subjected firstly to ageingA for 6 hours at 175° C.+13 hours at 190° C., corresponding to a T78temper with 55 hours of equivalent time at 175° C. and secondly ageing Bfor 6 hours at 175° C.+6 hours at 190°, in a slightly over-aged temperwith an equivalent time at 175° C. of 31 hours, making it possibleexacerbate the sensitivity to intercrystalline corrosion. The yieldstrength (in MPa) and the elongation (in %) in the TL direction in theT4 temper and the yield strength in the TL direction in the T78 temperafter ageing A (Table 17) were measured, along with the sensitivity tointercrystalline corrosion for the ageing steps A and B, indicating thedepth of the corrosion (in μm) and extent of the corrosion as a % of thesurface affected (Table 18). TABLE 17 Finishing R_(0.2) (T4-TL) R_(0.2)(T78-TL) A (T4-TL) None 191 337 24.6 3.2% traction 234 330 21.7 2.6%rolling 235 333 21.4 3.5% rolling 236 332 21.1 5.3% rolling 261 336 18.78.7% rolling 285 340 16.4

[0097] TABLE 18 Ageing A Ageing B IC Depth IC Depth Finishingsensitivity Extent sensitivity Extent None Yes 190 μm-10% Yes 150 μm-20%3.2% Slight  10 μm Yes 140 μm-1% traction 2.6% Slight  10 μm Yes 190μm-5% rolling 3.5% No — Yes 125 μm-10% rolling 5.3% No — Yes  25 μm-1%rolling 8.7% No — No — rolling

[0098] The correlation between the yield strength in the T4 temper andthe desensitization to intercrystalline corrosion in the T78 temper isagain observed. It is also observed that a high cold-working rate doesnot induce a degradation of the yield strength after ageing, as could beexpected, since the ageing kinetics is accelerated.

What is claimed is:
 1. A method for production of aircraft structuralcomponents using rolled, extruded or forged products made of an aluminumalloy, comprising the steps of: casting a blank with a compositionconsisting essentially of, in % by weight: Si: 0.7-1.3; Mg: 0.6-1.1; Cu:0.5-1.1; Mn: 0.3-0.8; Zn<1; Fe<0.30; Zr<0.20; Cr<0.25; other elements<0.05 each and <0.15 in total; remainder aluminum; hot, and possiblycold, transforming of said blank to obtain a product; solution heattreating the product between 540 and 570° C.; quenching the solutionheat treated product; forming and optionally welding the quenchedproduct to produce the structural component; ageing the structuralcomponent, in one or more stages, for a total equivalent time at 175° C.expressed in hours between (−160+57γ) and (−184+69γ), γ being Si+2Mg+2Cuin % by weight.
 2. Method according to claim 1, wherein the blank ishomogenized at a temperature between 540 and 570° C.
 3. Method accordingto claim 1, wherein the quenched product is subjected, before ageing, tocold-working resulting in a permanent elongation between 1 and 15%. 4.Method according to claim 3, wherein the cold working results in apermanent elongation between 2 and 10%.
 5. Method according to claim 1,wherein total equivalent time at 175° C., in hours is between (−150+57γ)and (−184+69γ).
 6. Method according to claim 1, wherein the product hasa composition, in % by weight: Si: 0.7-1.1; Mg: 0.6-0.9; Cu: 0.5-0.7;Mn: 0.3-0.8; Zr<0.2; Fe<0.2; Zn<0.5; Cr<0.25; Mg/Si<1; Si+2 Mg: 2.0-2.6;other elements <0.05 each and <0.15 in total; remainder alumnium. 7.Method according to claim 6, wherein Si+2Mg is between 2.3 and 2.6. 8.Method according to claim 6, wherein total equivalent ageing time at175° C. is between 40 and 65 hours.
 9. Method according to claim 3,additionally comprising assembly of sheets and profiles produced by themethod, wherein the profiles undergo, before said assembly and saidageing, an additional cold-working step in relation to the sheets, suchthat resistance to intercrystalline corrosion of the profiles is at thesame level as the sheets.
 10. Aircraft fuselage component, producedusing a rolled, extruded or forged product made of an aluminum alloyconsisting essentially of, in % by weight: Si: 0.7-1.1; Mg: 0.6-0.9; Cu:0.5-0.7; Mn: 0.3-0.8; Zr<0.2; Fe<0.2; Zn<0.5; Cr<0.25; Mg/Si<1; Si+2 Mg:2-2.6; other elements <0.05 each and <0.15 in total; remainder aluminum;subjected to a solution heat treatment, quenching, shaping and ageing ina T78 temper with a total equivalent time at 175° C. between 40 and 65hours.
 11. Fuselage component according to claim 10, wherein Si+2Mg isbetween 2.3 and 2.6.
 12. Fuselage component according to claim 10,having, in TL direction, a yield strength R_(0.2)>330 MPa, an ultimatetensile strength R_(m)>360 MPa and an elongation A>8%.
 13. Fuselagecomponent according to claim 10, having a plane strain fracturetoughness in T-L direction, with at least one of the properties: K_(R)(Δa=20 mm)>90 MPa{square root}m; K_(R) (Δa=40 mm)>115 MPa{square root}m;K_(c0)>80 MPa{square root}m; K_(c)>110 MPa{square root}m.
 14. Fuselagecomponent according to claim 10, having a plane strain fracturetoughness in L-T direction such that: K_(c0)>90 MPa{square root}m orK_(c)>130 MPa{square root}m.
 15. Fuselage component according to claim10, having a crack growth rate da/dn, measured in T-L direction forR=0.1, less than: 2 10⁻³ mm/cycle for ΔK=20 MPa{square root}m; 4 10⁻³mm/cycle for ΔK=25 MPa{square root}m; 8 10⁻³ mm/cycle for ΔK=30MPa{square root}m.
 16. Fuselage component produced using a rolled,extruded or forged product made of alloy consisting essentially of, in %by weight: Si: 0.7-1.1; Mg: 0.6-0.9; Cu: 0.5-0.7; Mn: 0.3-0.8; Zr<0.2;Fe<0.2; Zn<0.5; Cr<0.25; Mg/Si<1; Si+2 Mg: 2-2.6; other elements <0.05each and <0.15 in total; the remainder aluminum; subjected to a solutionheat treatment, quenching, shaping and ageing in T6 temper.
 17. Fuselagecomponent according to claim 16, having in TL direction, a yieldstrength R_(0.2)>350 MPa, an ultimate tensile strength R_(m)>380 MPa andan elongation A>6%.
 18. Fuselage component according to claim 16, havinga plane strain fracture toughness in T-L direction, with at least one ofthe properties: K_(R) (Δa=20 mm)>95 MPa{square root}m; K_(R) (Δa=40mm)>120 MPa{square root}m; K_(c0)>85 MPa{square root}m; K_(c)>115MPa{square root}m.
 19. Fuselage component according to claim 16, havinga plane strain fracture toughness in L-T direction such that: K_(c0)>100MPa{square root}m or K_(c)>150 MPa{square root}m.
 20. Fuselage componentaccording to claim 16, having a crack growth rate da/dn, measured in T-Ldirection for R=0.1, less than: 2 10⁻³ mm/cycle for ΔK=20 MPa{squareroot}m; 4 10⁻³ mm/cycle for ΔK=25 MPa{square root}m; 8 10⁻³ mm/cycle forΔK=30 MPa{square root}m.